1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine rotor blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine with rotor blades and stator blades that are exposed to a hot gas flow in order to convert combustion energy into mechanical energy. The turbine efficiency, and therefore the engine efficiency, can be increased by passing a higher temperature gas flow through the turbine, referred to as the turbine inlet temperature. The highest turbine inlet temperature is limited to both the material properties of the airfoils (both blades and vanes have airfoils) and the amount of cooling that can be produced in these airfoils.
FIG. 1 shows a prior art turbine rotor blade of U.S. Pat. No. 5,702,232 issued to Moore on Dec. 30, 1997 and entitled COOLED AIRFOILS FOR A GAS TURBINE ENGINE. This blade uses near wall cooling in the airfoil mid-chord section that is constructed with radial flow channels plus resupply holes in conjunction with film discharge cooling holes. In this design, the spanwise and chordwise cooling air flow control due to airfoil external hot gas temperature and pressure variation is difficult to achieve. In addition, a single radial flow channel is not the best method of utilizing cooling air because this results in a low convective cooling effectiveness. Also, the dimension for the airfoil external wall has to meet the investment casting requirements.